Hall Thruster Development for Japanese Space Propulsion Programs*

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Trans. Japan Soc. Aero. Space Sci. Vol. 60, No. 5, pp. 320–326, 2017

Yushi HAMADA,1) Junhwi BAK,1) Rei KAWASHIMA,1) Hiroyuki KOIZUMI,1) Kimiya KOMURASAKI,1) Naoji YAMAMOTO,2)† Yusuke EGAWA,2) Ikkoh FUNAKI,3) Shigeyasu IIHARA,4) Shinatora CHO,5) Kenichi KUBOTA,5) Hiroki WATANABE,6) Kenji FUCHIGAMI,7) Yosuke TASHIRO,4) Yuya TAKAHATA,8) Tetsuo KAKUMA,8) Yusuke FURUKUBO,8) and Hirokazu TAHARA8)

1)Department of Aeronautics and Astronautics, The University of Tokyo, Tokyo 113–8656, Japan 2)Department of Advanced Energy Engineering Science, Kyushu University, Kasuga, Fukuoka 816–8580, Japan 3)Institute of Space and Astronautical Science, JAXA, Sagamihara, Kanagawa 252–5210, Japan 4)IHI Aerospace Co., Ltd., Tomioka, Gunma 370–2398, Japan 5)Research and Development Directorate, JAXA, Chofu, Tokyo 182–8522, Japan 6)Department of Aerospace Engineering, Tokyo Metropolitan University, Tokyo 191–0065, Japan 7)IHI Corporation, Yokohama, Kanagawa 235–8501, Japan 8)Department of Aeronautics and Astronautics, Osaka Institute of Technology, Osaka 535–8585, Japan

Three different types of high power Hall thrusters—anode layer type, magnetic layer type with high specific impulse, and magnetic layer type with dual mode operation (high thrust mode and high specific impulse mode)—have been developed, and the thrust performance of each thruster has been evaluated. The thrust of the anode layer type thruster is in the range of 19–219 mN, with power in the range of 325–4500 W. The thrust of the high specific impulse magnetic layer type thruster was 102 mN, with specific impulse of 3300 s. The thrust of the bimodal operation magnetic layer thruster was 385 mN with specific impulse of 1200 s, and 300 mN with specific impulse of 2330 s. The performance of these thrusters demonstrates that the Japanese electric propulsion community has the capability to develop a thruster for commercial use.

Key Words: Hall Thruster, In-space Propulsion

Nomenclature and construction of the Space Solar Power System. US, European, and Japanese research organizations and compa- F: thrust nies are competing in the development of high power electric g: gravitational acceleration propulsion.3–5) The most mature EP system, the ion engine Id: discharge current system, has shown good performance at low power levels, Isp: specific impulse from several hundred W to kW; the successes of the “Deep m_: mass flow rate Space 1”6) and the asteroid explorer “Hayabusa”7) have Pcoil: power consumption at coils shown the superiority of the ion engine system. Hall thrusters Pcathode: power consumption at cathode are also promising, and they have shown competitive per- T/P: thrust to power ratio formance against ion thruster systems8,9) and superior per- 10,11) Vd: discharge voltage formance at higher power levels. Traditionally, thrust- t: thrust efficiency to-power ratio, specific impulse, and hours of operation for Subscripts Hall thrusters are around 50–60 mN/kW, 1500–2000 s and a: anode 6000–10000 h, respectively. The characteristics of the Hall c: cathode thruster are its high thrust density and wide power range operation, with acceptable efficiency. The use of Hall thrusters 1. Introduction in satellites yields lower trip times and lower system gross mass than the use of ion thrusters.12) High power electric propulsion (EP) is increasingly in de- There are two major types of Hall thrusters; the magnetic mand as a main propulsion system for planetary exploration layer type and the anode layer type.13,14) Many satellites with missions1) and all-electric satellites.2) It would also be appro- the magnetic layer type Hall thrusters have been launched priate for space cargo missions for the Manned Mars Mission and operated. In 2003, the SMART-1 mission presented by ESA adopted a PPS-1350, which was operated successfully. Total impulse was 1.1 million N0s and delta-V was 3.9 © 2017 The Japan Society for Aeronautical and Space Sciences + / 15) ff Presented at the 8th Asian Joint Conference on Propulsion and Power, km s. A Hall thruster named Hall E ect Rocket with Mag- March 16–19, 2016, Takamatsu, Kagawa, Japan, and the 60th Space netic Shielding (HERMeS) is currently under development Sciences and Technology Conference, September 6–9, 2016, Hakodate, for the US Solar Electric Propulsion demonstration mis- Hokkaido. sion.16) This is a 12.5 kW, 3000 s specific impulse thruster Received 26 August 2016; final revision received 24 December 2016; “ ”17) accepted for publication 17 March 2017. with magnetic shielding technology. Magnetic shielding †Corresponding author, [email protected] is expected to extend the life expectancy of the Hall thruster.

320 Trans. Japan Soc. Aero. Space Sci., Vol. 60, No. 5, 2017

The other type of Hall thruster, the anode layer type, was de- Outer coil veloped in Russia as a Thruster with Anode Layer (TAL). Trim coil The D-55 is the most famous Russian anode layer type Hall Anode thruster, but it has less operational time than the magnetic layer type Hall thrusters.18) In Japan, there have been many studies on Hall thrusters since the late 80’s,19) as reported by Kaufman.8) These stud- Inner coil ies have focused on the improvement of thrust performance in both single stage and two stage Hall thrusters (Periplasma- (a) (b) tron ion source,20) microwave discharge ion source21)), understanding of ion production/acceleration/loss processes, Fig. 1. RAIJIN94 Hall thruster, (a) Photo, (b) Magnetic coil position. and understanding of the plasma-wall interaction mechan- ism. In addition to experimental studies, since the 1990’s, numerical simulation codes have been developed to study the physics behind the Hall thruster.22–26) These studies have demonstrated that numerical simulation is a powerful tool for clarification of the physics in the system as well as for the development of Hall thrusters. Based on this experience, several Hall thrusters have recently been developed at the University of Tokyo,27,28) Osaka University/Osaka Institute of Technology (OIT),29,30) Tohoku University,31) Nagoya University,32) Gifu University,33) Kyushu University,34) Tokyo Metropoli- tan University (TMU),35) and The Institute of Space and As- tronautical Science (ISAS) in the Japan Aerospace Explora- 36) tion Agency (JAXA). This work has not been confined to (a) (b) universities and JAXA, industry has also been involved, in collaboration with universities; the IHI Corporation and Fig. 2. Schematic of SPT-type Hall thruster THT-VI, (a) Overview, (b) Osaka University37) developed a 1 kW class Hall thruster, Cross section. and Mitsubishi Electric Corporation38) developed a 250 mN class Hall thruster. tion. The thruster has a hollow annular anode, which consists The understanding developed through this work, and the of two cylindrical rings, with a propellant gas fed through successful demonstration of an all-electric propulsion satel- them. The gap between the tip of the anode and the exit of lite by Boeing,2) have inspired the development of high the acceleration channel is fixed at 3 mm. A hollow cathode power electric propulsion in Japan; three 5 kW class Hall (Veeco, HECS) is used as an electron source. Tests were con- thrusters are currently under development. One is a ducted in an ISAS/JAXA ion engine endurance test vacuum 2–6 kW class dual mode operation Hall thruster developed chamber,40) 2 m diameter by 5 m length, evacuated by four by JAXA, IHI, IHI Aerospace Engineering Corporation cryogenic pumps (44,000 l/s for xenon), with the pressure (IA), and TMU. (2–6 kW is considered to be the most useful kept below 6:6 103 Pa (for xenon) during thruster opera- operational range for a variety of missions.) The second is a tion, with total mass flow rates of 13.5 mg/s (140 sccm). 5 kW class anode layer type Hall thruster developed as part Pressure was measured using an ionization gauge, which of the Robust Anode-layer Intelligent Thruster for the Japa- was positioned close to the top of the thruster behind a nese IN-space propulsion system (RAIJIN) project, com- shroud. The chamber baseline pressure was below prised of nine universities and JAXA.39) The third is a mag- 1 105 Pa. High-purity (99.999%) xenon gas was used as netic layer type Hall thruster with high specific impulse the propellant. Thrust measurement was performed using a developed at OIT. In the present paper, we report on the cur- dual pendulum thrust stand.41) The uncertainty of the thrust rent status of these thrusters. was estimated as 3% at the thrust level of 200 mN, due to the friction of the calibration system, thermal drift effect, 2. Experimental Setup and the magnetic interference between the thruster and the thrust stand/vacuum facility. Figure 1 shows the 5 kW class anode layer type Hall The THT-VI magnetic layer type Hall thruster for high thruster, RAIJIN94, developed as part of the RAIJIN project. specific impulse, which was developed at Osaka Institute The inner and outer diameters of the acceleration channel are of Technology, is shown in Fig. 2. The outer diameter of 60 mm and 94 mm, respectively. An inner solenoid coil and the discharge channel is 100 mm and the inner diameter is four outer solenoid coils create a predominantly radial mag- 56 mm, that is, the channel is 22 mm wide and 40 mm long. netic field in the acceleration channel. There is a trim coil, These channel dimensions are the same as those of the Rus- which can change the shape of the magnetic field configura- sian SPT-100.42) The discharge channel wall is made of bor-

250

200

150

Thrust, mN 100 2.9 mg/s (1) (2) 4.9 mg/s 50 6.8 mg/s Fig. 3. Photo of breadboard models, (1) BBM1a, (2) BBM1b. 9.8 mg/s

0 0 1000 2000 3000 4000 5000 on nitride (BN). A hollow cathode (Veeco, HC-252) is used Power consumption, W as the electron source. Tests are conducted in the ISAS/ JAXA ion engine endurance test vacuum chamber with xe- Fig. 4. Thrust vs. power consumption for various mass flow rates for the non as the propellant. The thruster mass flow rate is fixed RAIJIN94 thruster. at 3.0 mg/s. The hollow cathode mass flow rate is 0.1 mg/s F2 except in high discharge voltage operation; it changes to ¼ P ð2Þ / – t 0.2 mg s when the discharge voltage is 950 1000 V for sta- 2ðÞm_ a þ m_ c VdId þ Pcoil þ Pcathode ble operation. The inner coil current, outer coil current and back coil current are fixed at 0.3 A, 0.3 A, and 0.9 A, respec- Figure 4 shows the thrust of the RAIJIN94 5 kW class tively. anode layer thruster versus input power. The dependency Dual mode operation magnetic layer type Hall thruster of the thrust on incident power shows some inconsistency, breadboard models were developed by JAXA, IA, IHI/IA, as the thrust was measured for various mass flow rates and and TMU. For the requirements of the both high-thrust and various magnetic field configurations. The thrust is in the high-Isp modes, we intend to put the ion production region range of 19–219 mN for power in the range of 325– at the appropriate downstream position. The effective diam- 4500 W. The lower limit is imposed by low efficiency and eters of the channels are 100 mm for BBM1a and 140 mm for the upper limit is imposed by a combination of the cathode BBM1b, as shown in Fig. 3. A larger model, BBM2, 170- capacity, instability, and overheating. As in conventional mm effective diameter was also fabricated and tested. The Hall thrusters, the thrust and power consumption are almost design of the thruster head was based on pre-design numer- proportional to mass flow rate. The thrust is almost propor- ical calculations; not only the geometric and magnetic field tional to the square root of the discharge voltage, and power configurations, but also plasma production and loss mecha- consumption is almost proportional to the discharge voltage; nisms were quantitatively estimated prior to fabrication.43) thrust is proportional to the square root of power consump- The use of numerical simulation in the preliminary design tion, if the mass flow rate is fixed. phase allowed optimization of the configuration and drasti- Figure 5 shows the thrust versus coil current at a discharge cally shortened the development time. voltage of 300 V, mass flow rate of 4.9 mg/s, and trim coil Performance tests were conducted in the IHI vacuum current of 0 A. The thrust depends strongly on the magnetic chamber (2 m diameter and 3 m length) and the vacuum field configuration, since the ion production and acceleration chamber at the Georgia Institute of Technology (5 m diame- depend strongly on magnetic field configuration.20) The ter and 9 m length).44) In the experiment, the pressure inside thrust at the inner coil current of 0.5 A and the outer coil cur- the vacuum chamber was kept below 4 103 Pa, with a few rent of 0.84 A goes to a minimum value, 71 mN, in these exceptions. In the experiment, xenon mass flow rates were measurements. The thrust reaches maximum, 91 mN, at inner changed from 5 to 30 mg/s for the anode, and the cathode coil current of 0.3 A and outer coil current of 0.5 A. There is a mass fraction was kept to 10% of that of the anode. Discharge 20 mN difference, or about 25% of the average thrust, be- voltages were set at 150 to 800 V. tween the two different magnetic field configurations, even when the mass flow rate and discharge voltage are fixed. 3. Results and Discussion Figure 6 shows the relation between trim coil current and thrust at the inner coil current of 0.4 A, the outer coil current For evaluation of the Hall thruster performance, specific of 0.4 A, discharge voltage of 300 V, and mass flow rate of impulse, Isp, and thrust efficiency, t, are defined as 4.9 mg/s. The increase in the trim coil pushes the lines of F force toward the downstream region, as shown in Fig. 7. I ¼ ð1Þ sp This will also push the plasma generation and acceleration ðm_ a þ m_ cÞg region toward the downstream end of the field. This would

0.8 Thrust, mN 95 0.7 90 0.6 85

0.5 80 (a) (b) (c) Inner coil current, A 0.4 Fig. 7. Calculated magnetic field configuration for three values of trim coil 75 current (inner coil current of 0.4 A and outer coil current of 0.4 A), (a) Trim coil current of ¹1 A, (b) Trim coil current of 1 A, (c) Trim coil current of 4 A (calculated using Magnum4.0, Field Precision LLC). 0.3 70 0.3 0.4 0.5 0.6 0.7 0.8 Outer coil current, A 120 fi Fig. 5. Thrust versus coil con gurations. 100

80 5 60 4

Thrust, mN 40

3 20

2 0 0 500 1000 1500 2000 2500 3000 3500 4000 Power consumption, W Discharge current, A 1 (a) 900 0.7 0.6

80 0.5 0.4 0.3

Thrust, mN Thrust, 70 0.2 Thrust efficiency 0.1 60 0.0 -3 -2 -1 0 1 2 3 4 5 6 200 300 400 500 600 700 800 900 1000 1100 Trim coil current, A Discharge voltage, V (b) Fig. 6. Thrust and discharge current for various trim coil configurations. Fig. 8. THT-VI thruster thrust performance at mass flow rate of 3 mg/s, inner coil current of 0.3 A, outer coil current of 0.3 A, and trim coil current tend to extend the lifetime of the thruster, though the thrust of 0.9 A, (a) Thrust vs. power consumption, (b) Thrust efficiency vs. dis- and discharge current might be decreased due to degradation charge voltage. of propellant utilization. This result shows that with increase in trim coil current, the thrust and discharge current decrease, The magnetic coil ratio and magnitude are optimized at the as mentioned above, and the thrust achieves a maximum val- discharge voltage of 750 V and mass flow of 3.0 mg/s. The ue of 82 mN at the trim coil of ¹1 A. The thrust reaches a magnetic coil ratio and magnitude are fixed to demonstrate maximum at trim coil of ¹1 A, because with a decrease in the dependency of current and thrust on discharge voltage trim coil current, the lines of force are attracted into the hol- in a fixed magnetic field configuration. Thrust is almost pro- low anode, and the ion production region moves upstream. portional to the square root of the input power, as in conven- This improves propellant utilization and thrust increases. It tional Hall thrusters, and thrust of 100 mN was achieved at should be noted that excessive trim current does not affect discharge voltage of 850 V. THT-VI is a high specific im- the plasma generation region due to the anode geometry. pulse aimed Hall thruster and, indeed, the specific impulse These results show that magnetic field configuration was 3:3 103 s at discharge voltage of 850 V, with good greatly affects both thrust and power consumption; optimiza- thrust efficiency of 0.60. This is a sufficiently high specific tion of the magnetic field is crucial, not only to the lifetime of impulse to perform planetary exploration missions. The the thruster17) but also to thruster performance. THT-VI shows stable operation below a discharge voltage Figure 8 shows the relation between power consumption of 900 V, and it shows unstable operation above 900 V, ow- and thrust of the THT-VI at a mass flow rate of 3.0 mg/s. ing to overheating of the thruster body; the outer acceleration

500 90 η η=0.6 η=0.4 =0.5 80

400 70 BBM1a-6mg/s BBM1a-10mg/s 60 BBM1b-6mg/s 300 50 BBM1b-12mg/s BBM1b-20mg/s 40 BBM1b-30mg/s

Thrust,mN 200 BBM2-6mg/s 30 THT VI BBM2-12mg/s 20 BBM2-20mg/s RAIJIN 94 BBM2-30mg/s to power ratio, mN/kW Thrust 10 IHI/JAXA/TMU_BBM2 100 0 0 500 1000 1500 2000 2500 3000 3500 0 Specific impulse, s 0 2000 4000 6000 8000 1 10441.2 10 Fig. 10. Thrust to power ratio versus speciﬁc impulse. Power Consumption, W

Fig. 9. Thrust characteristics of breadboard models (BBM1a, BBM1b and Table 1. Comparison of thrusters.46) BBM2; mass flow rate is for anode mass flow rate).45) Thrusters Power, kW Efficiency Specific impulse, s RAIJIN94 4.5 0.53 2200 THT-VI 2.7 0.60 3300 channel wall was observed to be red hot. Overheating could IHI/JAXA/IA/TMU 4 0.40 1050 be overcome by the adoption of a large radiation shield, in- BBM2 7 0.54 2300 crease in the thruster size, and/or the adoption of a pyrolytic SPT-140 4.5 0.55 1800 carbon anode. Maximum efficiency is 0.60 at discharge volt- BPT-4000 4.5 0.57 2000 age of 750 V; thrust performance would be improved if the magnetic field configuration (inner coil current, outer coil current and trim coil current) were optimized for the dis- GTO to GEO or drag compensation at low altitude orbit. charge voltage. Figure 10 shows T/P for three thrusters versus specific im- The thrust of the breadboard models developed at IHI/ pulse. The T/P is in inverse proportion to the specific im- JAXA/IA/TMU, the BBM1a, BBM1b and BBM2 thrusters, pulse, if the thrust efficiency is constant. The BBM2 has a is plotted against input power in Fig. 9. The magnetic field high T/P with specific impulse of 1000–2000 s and the max- configuration was optimized for each condition (each mass imum T/P exceeds 70 mN/kW. The BBM2 is a remarkable flow and each discharge voltage). These thrusters are de- thruster, since it achieved a high thrust efficiency of 0.4 at a signed for dual mode operation, that is, high thrust mode low specific impulse of 1050 s, and higher efficiency at high- for quick orbital transfer from GTO to GEO, and high specif- er specific impulses. ic impulse mode for north-south station keeping (NSSK). Conventional Hall thrusters have a low T/P with specific The target thrust of the high thrust mode is 320–430 mN, impulse of less than 1000 s, as a results of low thrust effi- with specific impulse of 1180 s or higher. The second target ciency. this is due to low propellant utilization for discharge thrust is 140 mN at power consumption of 3 kW with specific voltages below 150 V. The THT-VI also has a high T/P ratio, impulse of 2500 s or higher. BBM1b demonstrates good bi- considering its specific impulse of 1800–3300 s. This is due modal operation; thrust of 354 mN with specific impulse of to high thrust efficiency (0.5–0.6). RAIJIN94 also shows 1100 s at power consumption of 5140 W and mass flow rate high T/P, with specific impulse of 1200 s to 2000 s. of 30 mg/s, and 130 mN with specific impulse of 2350 s at These thrusters have a wide variety of T/P (from 30 mN mass flow rate of 5 mg/s and power consumption of to 60 mN) and specific impulse (from 1800 s to 3300 s), 3500 W. BBM2 also demonstrated dual mode operation; and they demonstrate capable performance for practical ap- thrust of 380 mN with specific impulse of 1200 s at mass flow plications, as shown in Table 1. Lifetime assessment of the rate of 30 mg/s, and thrust of 140 mN with specific impulse thruster head and development of a neutralizer will be neces- of 2050 s at mass flow rate of 6 mg/s and 300 mN with spe- sary next steps. cific impulse of 2330 s at mass flow rate of 12 mg/s. The upper limit of input power is defined by thermal issues and 4. Conclusion facility restrictions, the thrust range is limited to 136– 473 mN for BBM2 in the power range of 1500 W to Several 5 kW class Hall thrusters, RAIJIN94, THT-VI and 6970 W. The thrust efficiency is in the range of 0.4 to 0.6, BBM1a, BBM1b and BBM2 have been developed at UT/ which is quite good as compared to the other Hall thrust- KU, OIT, and JAXA/IA/IHI/TMU, and the thrust perform- ers.14–17) ance evaluated. The specific impulse was found to be in the Thrust-to-power ratio (T/P) is a good indicator of thruster range of 1000 s to 3300 s with sufficient thrust efficiency performance--large T/P is required for orbit transfer from (0.4–0.6) and thrust to power ratio (30–78 mN/kW). The

©2017 JSASS 324 Trans. Japan Soc. Aero. Space Sci., Vol. 60, No. 5, 2017 thrust to power ratio of BBM2 was 78 mN/kW, with specific Applications (ICRERA), 2012. 13) Choueiri, E. Y.: Fundamental Difference between the Two Hall impulse of 1050 s; this is a remarkable thrust to power ratio – ffi Thruster Variants, Phys. Plasmas, 8 (2001), pp. 5025 5033. and it would be su cient to shorten the transition from GTO 14) Zhurin, V. V., Kaufman, H. R., and Robinson, R. S.: Physics of Closed to GEO. The specific impulse of THT-VI achieved 3300 s Drift Thrusters, Plasma Sources Sci. Technol., 8 (1999), pp. R1–R20. with a high thrust efficiency of 0.60; this performance rivals 15) Racca, G. D.: SMART-1 from Conception to Moon Impact, J. Propul. – / that of other high specific impulse Hall thrusters. RAIJIN94 Power, 25 (2009), pp. 993 1002, doi:10.2514 1.36278 ffi – 16) Hofer, R., Kamhawi, H., Herman, D., Polk, J., Snyder, J. S., shows good performance, and thrust e ciency of 0.5 0.6 Mikellides, J., Huang, W., Myers, J., Yim, J., Williams, G., Ortega, with specific impulse of 1200–2200 s. Hall thruster develop- A. L., Jorns, B., Sekerak, M., Griffiths, C., Shastry, R., Haag, T., ment in Japan has been highly successful, and has led to the Verhey, T., Gilliam, B., Katz, I., Goebel, D., Anderson, J. R., Gilland, current state-of-the-art. The stage is set for technological and J., and Clayman, L.: Development Approach and Status of the 12.5 kW HERMeS Hall Thruster for the Solar Electric Propulsion Technology application breakthroughs in the near future. Demonstration Mission, IEPC 2015-186, 2015. 17) Mikellides, I. G., Katz, I., Hofer, R. R., Goebel, D. M., Grys, K., and Acknowledgments Mathers, A.: Magnetic Shielding of the Channel Walls in a Hall Plas- ma Accelerator, Phys. Plasmas, 18 (2011), 033501, doi:10.1063/ 1.3551583 The authors gratefully acknowledge support from JAXA, IA and 18) Oleson, S. R. and Sankovic, J. M.: Advanced Hall Electric Propulsion IHI. Research results were obtained using the ion engine endurance for Future In-Space Transportation, NASA/TM-2001-210676, 2001. test facility at ISAS, JAXA. The contributions of Prof. Hitoshi 19) Yamagiwa, Y. and Kuriki, K.: Performance of Double-Stage-Dis- Kuninaka, Prof. Kazutaka Nishiyama, Prof. Satoshi Hosoda, Prof. charge Hall Ion Thruster, J. Propul. Power, 7 (1991), pp. 65–70, Ryudo Tsukizaki, and their staff are greatly appreciated. For the doi:10.2514/3.23295 experiment at GIT, the contributions of Prof. Mitchell Walker and 20) Komurasaki, K. and Arakawa, Y.: Hall Thruster Performance and Plas- his staff are much appreciated. ma Acceleration Processes, J. Jpn. Soc. Aeronaut. Space Sci., 40 (1992), pp. 568–575 (in Japanese). 21) Kuwano, H., Kuninaka, H., and Nakashima, H.: Plasma Characteristics in the Acceleration Channel of a Microwave Discharge Hall Thruster References and Relationships between Thruster Performance and Acceleration Channel Length, J. Jpn. Soc. Aeronaut. Space Sci., 55 (2007), 1) Juno/NASA, http://www.nasa.gov/mission_pages/juno/main/index.html pp. 188–194 (in Japanese). (accessed 7 June 2016). 22) Komurasaki, K. and Arakawa, Y.: Two-dimensional Numerical Model 2) Boeing: World’s First All-Electric Propulsion Satellite Begins Opera- of Plasma Flow in a Hall Thruster, J. Propul. Power, 11 (1995), tions, http://boeing.mediaroom.com/2015-09-10-Boeing-World-s- pp. 1317–1323, doi:10.2514/3.23974 First-All-Electric-Propulsion-Satellite-Begins-Operations (accessed 7 23) Hirakawa, M.: Particle Simulation of Plasma Phenomena in Hall June 2016). Thrusters, IEPC 95-164, 1995. 3) Florenz, R. E.: The X3 100-kW Class Nested-Channel Hall Thruster: 24) Tahara, H., Shirasaki, A., and Martinez-Sanchez, M.: One-Dimension- Motivation, Implementation, and Initial Performance, Ph.D. Thesis, al Calculation of Hall Thruster Flowfields, J. Jpn. Soc. Aeronaut. University of Michigan, 2013. Space Sci., 51 (2003), pp. 1–9 (in Japanese). 4) Soulas, G., Haag, T., Herman, D., Huang, W., Kamhawi, H., and 25) Yokota, S., Takahashi, D., Cho, S., Kaneko, R., Hosoda, M., Shastry, R.: Performance Test Results of the NASA-457M v2 Hall Komurasaki, K., and Arakawa, Y.: Magnetic Topology to Stabilize Thruster, AIAA 2012-3940, 2012. Ionization Oscillation in Anode-layer-type Hall Thruster, Trans. 5) Semenkin, A. V., Zakharenkov, L. E., and Soldukhin, A. E.: Feasibil- JSASS Aerospace Technology Japan, 10, ists28 (2012), pp. Pb_31– ity of High Power Multi-Mode EPS Development Based on the Pb_35. Thruster with Anode Layer, IEPC 2011-064, 2011. 26) Cho, S., Yokota, S., Kaneko, R., Komurasaki, K., and Arakawa, Y.: 6) Marcucci, M. G. and Polk, J. E.: NSTAR Xenon Ion Thruster on Deep Channel Wall Erosion Modeling of a SPT-Type Hall Thruster, Trans. Space 1: Ground and Flight Tests, Rev. Sci. Instrum., 71 (2000), JSASS Aerospace Technology Japan, 10, ists28 (2012), pp. Pb_25– pp. 1389–1400, http:// dx.doi.org/10.1063/1.1150468 Pb_30. 7) Kuninaka, H., Nishiyama, K., Funaki, I., Yamada, T., Shimizu, Y., and 27) Komurasaki, K. and Arakawa, Y.: Hall-Current Ion Thruster Perform- Kawaguchi, J.: Deep Space Flight of Microwave Discharge Ion En- ance, J. Propul. Power, 8 (1992), pp. 1212–1216. gines Onboard HAYABUSA, J. Space Technol. Sci., 22 (2006), 28) Yamamoto, N., Nakagawa, T., Komurasaki, K., and Arakawa, Y.: Op- pp. 1–10. erating Characteristics of an Anode Layer Type Hall Thruster, J. Jpn. 8) Kaufman, H. R.: Technology of Closed-Drift Thrusters, AIAA J., 23 Soc. Aeronaut. Space Sci., 51 (2003), pp. 492–497 (in Japanese). (1985), pp. 78–86. 29) Yuge, S., Shirasaki, A., and Tahara, H.: Basic Operational Character- 9) Kim, V.: Main Physical Feature and Processes Determining the Per- istics and Thrust Performance of Low Power Hall Thrusters, J. Jpn. formance of Stationary Plasma Thrusters, J. Propul. Power, 14 Soc. Aeronaut. Space Sci., 55 (2007), pp. 8–16 (in Japanese). (1998), pp. 736–743. 30) Shirasaki, A. and Tahara, H.: Operational Characteristics and Plasma 10) Kamhawi, H., Huang, W., Haag, T., Shastry, R., Thomas, R., Yim, J., Measurements in Cylindrical Hall Thrusters, J. Appl. Phys., 101 Herman, D., Williams, G., Myers, J., Hofer, R., Mikellides, I., Sekerak, (2007), 073307, http:// dx.doi.org/10.1063/1.2720093 M., and Polk, J.: Performance and Facility Background Pressure Char- 31) Ando, A., Tashiro, M., Hitomi, K., Hattori, K., and Inutake, M.: Beam acterization Tests of NASA’s 12.5-kW Hall Effect Rocket with Mag- Extraction from a Hall-type Ion Accelerator, Rev. Sci. Instrum., 79 netic Shielding Thruster, IEPC 2015-07, July, 2015. (2008), 02B705. 11) Sengupta, A., Marrese-Reading, C., Semenkin, A. V., Zakharenkov, 32) Komurasaki, K. and Kusamoto, D.: Optical Measurement of Plasma L., Tverdokhlebov, S., Tverdokhlebov, O., Polzin, K., Markusic, T., Oscillations in a Hall Thruster, Trans. Jpn. Soc. Aeronaut. Space Cappelli, M., Scharfe, D., Boyd, I., Keidar, M., Yalin, A., and Surla, Sci., 40 (1999), pp. 203–208. V.: Summary of the VHITAL Thruster Technology Demonstration 33) Miyasaka, T., Shibata, Y., Asato, K., and Segawa, K.: Investigation of Program: A Two-Stage Bismuth-Fed Very High Specific Impulse Acceleration Channel Processes in Hall Thrusters by Particles Simula- TAL, IEPC 2007-005, 2007. tions, Trans. JSASS Space Technology Japan, 7, ists26 (2009), 12) Ito, Y., Nakano, M., Schonherr, T., Cho, S., Komurasaki, K., and pp. Pb_83–Pb_88. Koizumi, H.: In-Space Transportation of a Solar Power Satellite Using 34) Yamamoto, N., Tao, L., Rubin, B., Williams, J. D., and Yalin, A. P.: a Hall Thruster Propulsion System, Renewable Energy Research and Sputter Erosion Sensor for Anode Layer-Type Hall Thrusters Using

Cavity Ring-Down Spectroscopy, J. Propul. Power, 26 (2010), and Science, 1998, pp. 318–323. pp. 142–148. 41) Nagao, N., Yokota, S., Komurasaki, K., and Arakawa, Y.: Develop- 35) Watanabe, H., Ichimura, M., and Takegahara, H.: Operating Charac- ment of a 2D Dual Pendulum Thrust Stand for Hall Thrusters, Rev. teristics of Hall Thruster with Radio Frequency Plasma Cathode, J. Sci. Instrum., 76 (2007), 115108. Jpn. Soc. Aeronaut. Space Sci., 64 (2016), pp. 171–181 (in Japanese), 42) Garner, C. E., Brophy, J. R., Polk, J. E., and Pless, L. C.: A 5,730-Hr doi: 10.2322/jjsass.64.171 Cyclic Endurance Test of the SPT-100, AIAA 95-2667, July 1995. 36) Molina-Morales, P., Kuninaka, H., Toki, K., and Arakawa, Y.: Pre- 43) Cho, S., Watanabe, H., Kubota, K., Iihara, S., Fuchigami, K., Uematsu, liminary Study of an ECR Discharge Hall Thruster, IEPC 2001-069, K., and Funaki, I.: Study of Electron Transport in a Hall Thruster by 2001. Axial–radial Fully Kinetic Particle Simulation, Phys. Plasmas, 22 37) Kitano, T., Fujioka, T., Shirasaki, A., Goto, D., Tahara, H., Yasui, T., (2015), 103523. Yoshikawa, T., Fuchigami, K., Iinoya, F., and Ueno, F.: Research and 44) Vacuum Test Facility-2, High-Power Electric Propulsion Laboratory at Development of Low Power Hall Thrusters, 23rd International Sympo- the Georgia Institute of Technology, http://mwalker.gatech.edu/hpepl/ sium on Space Technology and Science, Matsue, ISTS 2002-b-18, facilities/vtf2/ (accessed 7 June 2016). 2002. 45) Funaki, I., Iihara, S., Cho, S., Kubota, K., Watanabe, H., Fuchigami, 38) Osuga, H., Suzuki, K., Nakagawa, T., Ozaki, T., Tamida, T., and K., and Tashiro, Y.: Laboratory Testing of Hall Thrusters for All-elec- Matsui, K.: Performance of Power Processing Unit for 250 mN-class tric Propulsion Satellite and Deep Space Explorers, 52nd AIAA/SAE/ Hall Thruster, IEPC 2009-117, 2009. ASEE Joint Propulsion Conference, AIAA 2016-4942, Salt Lake City, 39) Yamamoto, N., Miyasaka, T., Komurasaki, K., Koizumi, H., Utah, July 2016. Schönherr, T., Tahara, H., Takegahara, H., Aoyagi, J., Nakano, M., 46) Oh, D. Y., Snyder, J. S., Goebel, D. M., Hofer, R. R., and Funaki, I., Watanabe, H., Ohkawa, Y., Kakami, A., Takao, Y., Yokota, Randolph, T. M.: Solar Electric Propulsion for Discovery-Class Mis- S., Ozaki, T., and Osuga, H.: Development of Robust Anode Layer In- sions, J. Spacecraft Rockets, 51 (2014), pp. 1822–1835. telligent Thruster for Japanese In Space Propulsion, IEPC 2013-244, 2013. Y. Ohkawa 40) Kuninaka, H., Funaki, I., Shimizu, Y., and Toki, K.: Endurance Test Facility and Test Status of Microwave Discharge Ion Thruster, Pro- Associate Editor ceedings of the 21st International Symposium on Space Technology